(1) Field of the Invention
The present invention relates to a turbine engine component having an improved scheme for cooling an airfoil portion.
(2) Prior Art
The overall cooling effectiveness is a measure used to determine the cooling characteristics of a particular design. The ideal non-achievable goal is unity, which implies that the metal temperature is the same as the coolant temperature inside an airfoil. The opposite can also occur when the cooling effectiveness is zero implying that the metal temperature is the same as the gas temperature. In that case, the blade material will certainly melt and burn away. In general, existing cooling technology allows the cooling effectiveness to be between 0.5 and 0.6. More advanced technology such as supercooling should be between 0.6 and 0.7. Microcircuit cooling as the most advanced cooling technology in existence today can be made to produce cooling effectiveness higher than 0.7.
FIG. 1 shows a durability map of cooling effectiveness (x-axis) vs. the film effectiveness (y-axis) for different lines of convective efficiency. Placed in the map is a point 10 related to a new advanced serpentine microcircuit shown in FIGS. 2a-2c. This serpentine microcircuit includes a pressure side serpentine circuit 20 and a suction side serpentine circuit 22 embedded in the airfoil walls 24 and 26.
The Table I below provides the operational parameters used to plot the design point in the durability map.
TABLE IOperational Parameters forserpentine microcircuitbeta2.898Tg2581 [F]Tc1365 [F]Tm2050 [F]Tm_bulk1709 [F]Phi_loc0.437Phi_bulk0.717Tco1640 [F]Tci1090 [F]eta_c_loc0.573eta_f0.296Total Cooling Flow3.503%WAE10.8Legend for Table IBeta = heat loadPhi_loc = local cooling effectivenessPhi_bulk = bulk cooling effectivenessEta_c_loc = local cooling efficiencyEta_f = film effectivenessTg = gas temperatureTc = coolant temperatureTm = metal temperatureTm_bulk = bulk metal temperatureTco = exit coolant temperatureTci = inlet coolant temperatureWAE = compressor engine flow, pps
It should be noted that the overall cooling effectiveness from the table is 0.717 for a film effectiveness of 0.296 and a convective efficiency (or ability to pick-up heat) of 0.573. Also note that the corresponding cooling flow for a turbine blade having this cooling microcircuit is 3.5% engine flow. FIG. 3 illustrates the cooling flow distribution for a turbine blade with the serpentine microcircuits of FIGS. 2a-2c embedded in the airfoils walls.
There are however field problems that can be addressed efficiently with peripheral microcircuit designs. One such field problem is illustrated in FIGS. 4A and 4B. In FIG. 4A, the streamlines of the gas path close to the external surface of the airfoil illustrate four different regions in which the gas flow changes direction or migration: a tip region, two mid-section regions, and a root region. In between the tip and the upper mid region, the flow transitions through a pseudo stagnation point(s). The momentum of the external gas seems to decelerate in such a way as to impose a local thermal load to the part. This manifests itself by regions where the propensity for erosion and oxidation increase in the airfoil surface. The superposition of FIG. 4B illustrates the local coincidence between the pseudo-stagnation region and the blade distress in the part surface. In the mid region, the upper and lower regions also converge onto one another, but even though the space between streamlines decreases, the flow seems to accelerate and there is no pseudo-stagnation regions. A mild manifestation of the same tip-to-mid phenomena seems to initiate in the transition region between the mid-to-root regions. It is therefore necessary to tailor the peripheral microcircuit in such a manner as to address these local high thermal load regions.